US11268441B2ActiveUtilityPatentIndex 71
Shaft bearing arrangement
Est. expiryDec 19, 2039(~13.5 yrs left)· nominal 20-yr term from priority
F02C 7/06F05D 2260/96F05D 2220/323F05D 2240/54F16C 2360/23Y02T50/60F01D 25/18F02C 7/36F01D 25/164F02C 3/06F05D 2240/50F01D 25/34F02K 3/06F01D 25/04F01D 25/16F05D 2240/52F01D 5/141F16C 19/18F02C 3/113F16C 19/54
71
PatentIndex Score
2
Cited by
61
References
20
Claims
Abstract
An engine core including a turbine, compressor, and core shaft connecting the turbine and compressor, the turbine being the lowest pressure turbine, the core shaft having a running speed range from 1500-6200 rpm, and the compressor being the lowest pressure compressor; a fan located upstream of the engine core; and a gearbox receiving an input from the core shaft and outputs drive to the fan. The engine core further includes three bearings arranged to support the core shaft, the three bearings including a forward bearing and two rearward bearings, the core shaft having a length between the forward and the rearmost rearward bearing ranging from 1800-2900 mm, and a minor span between two rearward bearings ranging from 250-350 mm, wherein there is no primary resonance of the core shaft between the forward and forwardmost rearward bearing within the running speed range of the core shaft.
Claims
exact text as granted — not AI-modifiedThe invention claimed is:
1. A gas turbine engine for an aircraft comprising:
an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and wherein the turbine is a lowest pressure turbine of the gas turbine engine, the core shaft has a running speed in a range from a lower bound of 1500 rpm to an upper bound of 6200 rpm, and the compressor is a lowest pressure compressor of the gas turbine engine;
a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and
a gearbox that is configured to receive an input from the core shaft and output drive to the fan so as to drive the fan at a lower rotational speed than a rotational speed of the core shaft, and wherein the engine core further comprises three bearings arranged to support the core shaft, the three bearings comprising a forward bearing and two rearward bearings, the core shaft having a length between the forward bearing and a rearmost bearing of the rearward bearings in a range from 1800 mm to 2900 mm, and a minor span, S, is defined as an axial distance between the two rearward bearings in a range from 250 mm to 350 mm, such that there is no primary resonance of the core shaft between the forward bearing and a forwardmost bearing of the rearward bearings within the running speed range of the core shaft.
2. The gas turbine engine of claim 1 , wherein the lower bound of 1500 rpm on the core shaft running speed is a minimum running speed under ground idle conditions and the upper bound of 6200 rpm on the core shaft running speed is the upper bound on maximum take-off running speed.
3. The gas turbine engine of claim 1 , wherein the fan has a fan diameter in a range from 330 cm to 380 cm; and the gearbox has a gear ratio in a range from 3.1 to 3.8.
4. The gas turbine engine of claim 1 , wherein the turbine has a turbine length defined between a leading edge of a most upstream blade of the turbine and a trailing edge of a most downstream blade of the turbine, and wherein a minor span to turbine length ratio of:
the
minor
span
the
turbine
length
is equal to or less than 1.05.
5. The gas turbine engine of claim 1 , wherein the turbine comprises four sets of turbine blades, and the two rearward bearings are located downstream of a trailing edge of a turbine blade of a third set of turbine blades from a front of the turbine, at a root of the turbine blade of the third set of turbine blades.
6. The gas turbine engine of claim 1 , wherein the turbine comprises a total of four sets of turbine blades.
7. The gas turbine engine of claim 1 , wherein the two rearward bearings are located downstream of a leading edge of a lowest pressure (most downstream) turbine blade of the turbine at a root of the lowest pressure turbine blade.
8. The gas turbine engine of claim 1 , wherein the length of the core shaft is in a range from 2000 to 2900 mm.
9. The gas turbine engine of claim 1 , wherein a length ratio of the minor span between the two rearward bearings to the core shaft length is equal to or less than 0.14.
10. The gas turbine engine of claim 1 , wherein the forwardmost bearing of the rearward bearings has a bearing stiffness in a range of 30 kN/mm to 100 kN/mm.
11. The gas turbine engine of claim 10 , wherein a stiffness ratio of the bearing stiffness at the forwardmost bearing of the rearward bearings to the distance between the two rearward bearings is in a range from 0.08 kN/mm 2 to 0.5 kN/mm 2 .
12. The gas turbine engine of claim 1 , wherein the lowest pressure turbine of the gas turbine engine has a lowest pressure set of blades, each blade of the lowest pressure set of blades having a mass, m, a radius at blade mid-height, r, and an angular velocity at cruise, ω; and wherein a first blade to bearing ratio of:
the
minor
span
S
mr
ω
2
(
for
a
blade
of
the
lowest
pressure
set
)
has a value in a range from 2.0×10 −6 kg −1 ·rad −2 ·s 2 to 7.5×10 −6 kg −1 ·rad −2 ·s 2 .
13. The gas turbine engine of claim 1 , wherein the lowest pressure turbine of the gas turbine engine has a lowest pressure set of blades, each blade of the lowest pressure set of blades having a mass, m, a radius at blade mid-height, r; and wherein a second blade to bearing ratio of:
the
minor
span
S
m
×
r
(
for
a
blade
of
the
lowest
pressure
set
)
has a value in a range from 0.8 kg −1 to 6.0 kg −1 .
14. The gas turbine engine of claim 1 , further comprising a tail bearing housing located rearward of the turbine and comprising two bearing discs, each bearing disc arranged to support one of the two rearward bearings.
15. The gas turbine engine of claim 14 , wherein the bearing discs are oriented at least substantially perpendicular to an axis of the gas turbine engine.
16. The gas turbine engine of claim 1 , wherein:
the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft;
the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and
the second turbine, the second compressor, and the second core shaft are arranged to rotate at a higher rotational speed than the rotational speed of the first core shaft.
17. A method of operation of a gas turbine engine for an aircraft comprising:
an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and wherein the turbine is a lowest pressure turbine of the gas turbine engine, and wherein the engine core further comprises three bearings arranged to support the core shaft, the three bearings comprising a forward bearing and two rearward bearings, the core shaft having a length between the forward bearing and a rearmost bearing of the rearward bearings in a range from 1800 mm to 2900 mm, and a minor span between the two rearward bearings in a range from 250 mm to 350 mm;
a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and
a gearbox arranged to receive an input from the core shaft and to output drive to the fan so as to drive the fan at a lower rotational speed than a rotational speed of the core shaft, the method comprising:
operating the gas turbine engine such that the core shaft has a running speed in a range from 1500 rpm to 6200 rpm, and wherein there is no primary resonance of the core shaft within the running speed range of the core shaft.
18. A method of designing a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and wherein the turbine is a lowest pressure turbine of the gas turbine engine, the core shaft has a running speed in a range from 1500 rpm to 6200 rpm, and the compressor is a lowest pressure compressor of the gas turbine engine;
a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and
a gearbox that is configured to receive an input from the core shaft and output drive to the fan so as to drive the fan at a lower rotational speed than a rotational speed of the core shaft, and wherein the engine core further comprises three bearings arranged to support the core shaft, the three bearings comprising a forward bearing and two rearward bearings, the core shaft having a length between the forward bearing and a rearmost bearing of the rearward bearings in a range from 1800 mm to 2900 mm, and the method comprising:
selecting positions for the forward bearing and a forwardmost bearing of the rearward bearings;
lengthening the core shaft rearward of the forwardmost bearing of the rearward bearings such that a minor span defined between the two rearward bearings is in a range from 250 mm to 350 mm, and there is no primary resonance of the core shaft between the forward bearing and the forwardmost bearing of the rearward bearings within the running speed range of the core shaft.
19. The method of claim 18 , wherein selecting positions for the forward bearing and the forwardmost bearing of the rearward bearings comprises locating the two rearward bearings downstream of a leading edge of a lowest pressure (most downstream) turbine blade of the turbine at a root of the lowest pressure turbine blade.
20. The method of claim 18 , wherein the turbine comprises four sets of turbine blades, and wherein selecting positions for the forward bearing and the forwardmost bearing of the rearward bearings comprises locating the two rearward bearings downstream of a trailing edge of a turbine blade of a third set of turbine blades from a front of the turbine, at a root of the turbine blade of the third set of turbine blades.Cited by (0)
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